Constant sun angle transfer orbit sequence and method using electric propulsion

ABSTRACT

An apparatus and method for translating a spacecraft (102, 108) from an injection orbit (114) about a central body (100) to synchronous orbit (122) in a time efficient manner. The spacecraft (102, 108) includes propulsion thrusters (50) which are fired in predetermined timing sequences controlled by a controller (64) in relation to the apogee (118) and perigee (120) of the injection orbit (114) and successive transfer orbits (114). During transfer orbit, the spacecraft&#39;s inertial attitude is adjusted to track sun movement. The spacecraft is injected into a particular injection orbit that offsets moments created by tracking sun movement to maintain stable transfer orbits.

RELATED APPLICATION

This Application is a continuation-in-part of application Ser. No.08/217,791 filed Mar. 25, 1994, now U.S. Pat. No. 5,595,360.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to an apparatus and method of translating aspacecraft from a particular injection orbit to a synchronous orbit in atime efficient manner while maintaining a constant sun angle to improvepower efficiency and allow spin stabilization control.

2. Description of the Related Art

In order to place a spacecraft into a final orbit, such as ageosynchronous orbit, about a central body, such as the earth, thespacecraft is first launched into an injection orbit by the spacecraftlaunch vehicle. From this injection orbit, the spacecraft is translatedthrough a series of orbits to the geosynchronous orbit. In order for thespacecraft to translate from its injection orbit to the geosynchronousorbit, propulsion thrusters fire to exert a force on the spacecraft andmove it through the transfer orbit.

There are a number of strategies for translating a spacecraft from itsinjection orbit to geosynchronous orbit. In a first strategy, a launchvehicle injects the spacecraft to an elliptical orbit having an apogeegreater than the geosynchronous orbit, defined as a supersynchronousorbit. Once the spacecraft has reached supersynchronous orbit,propulsion thrusters are fired when the spacecraft is in a predeterminedorientation and in proximity to apogee or perigee. Firing the propulsionthrusters at apogee to create thrust in the direction of orbitalvelocity raises perigee, and firing the propulsion thrusters at perigeeto create thrust in a direction opposite the orbital velocity lowersapogee. These apogee and perigee firings or bums translate thespacecraft from supersynchronous orbit to geosynchronous orbit. In asecond strategy, the spacecraft is injected into an elliptical orbithaving an apogee less than the geosynchronous, defined as asubsynchronous orbit. Once the spacecraft is in subsynchronous orbit,the propulsion thrusters are once again fired when the spacecraft inproximity to apogee and perigee and in a predetermined orientation.Firing at apogee to create thrust in the direction of orbital velocityraises perigee, and firing at perigee to create thrust in the directionof orbital velocity raises apogee. The apogee and perigee bums cause thespacecraft orbit to spiral out to the geosynchronous orbit. Such aspiraling-out mission using a specific type of thruster is described inMeserole, J. "Launch Costs to GEO Using Solar Powered Orbit TransferVehicles." American institute of Aeronautics and Astronautics (AIAA)Paper 93-2219, AIAA/SAE/ASME/ASEE 29th Joint Propulsion Conference andExhibit (Jun. 28-30, 1993).

European Publication No. 0 047 211 to Mortelette describes a transferorbit strategy in which the spacecraft is injected in a natural orbitwhere the apogee equals the semi-major axis of the desiredgeosynchronous orbit and the perigee is much smaller than the semi-majoraxis. The spacecraft applies thrust symmetrically about apogee forapproximately half the period of the orbit, to increase perigee to thesemi-major axis while leaving apogee unchanged. The spacecraft'sattitude changes constantly to remain aligned with the velocity vectorin order to provide maximum thrust. Because the launch vehicle injectsthe spacecraft into either a subsynchronous or supersynchronous orbit,the spacecraft must include its own propulsion system to effect atranslation from injection to geosynchronous orbit and to performorientation and other stationkeeping maneuvers. This raises severalconsiderations for selecting a particular injection orbit translationstrategy. Ideally, an injection orbit is selected so that the weight ofthe spacecraft without fuel, the dry weight, is maximized. The dryweight generally includes the weight of the instrumentation and theunderlying spacecraft structure for the instrumentation. Optimizing dryweight requires a trade-off between the capability of the launchvehicle, how high above the earth the spacecraft can be launched, andthe propulsion system of the spacecraft, the on-board thrusters and fuelcarried by the spacecraft to translate from injection orbit togeosynchronous orbit and perform stationkeeping maneuvers. Greaterinjection orbits, i.e., higher apogees, reduce the amount of propellantexpended by the spacecraft propulsion system to achieve geosynchronousorbit. On the other hand, the capability or payload capacity of thelaunch vehicle decreases with an increase in the targeted apogeealtitude, so that a more powerful launch vehicle is required to inject aspacecraft having the same mass to an injection orbit having a higherapogee. Thus, in order to optimize the weight of the spacecraft atarrival in geosynchronous orbit, defined as the beginning of life weight(BOL), there is a trade-off between the capability of the launch vehicleand the amount that the propulsion thrusters need to be fired. Ofcourse, the more that the propulsion thrusters are fired, morepropellant mass is required, leaving less mass allocated to usefulinstrumentation.

Further adding to the above considerations is that there are two typesof spacecraft propulsion thrusters, electric and chemical. Chemicalpropulsion thrusters provide the required thrust for translating thespacecraft from injection orbit to geosynchronous orbit and are capableof exerting a substantial force on the spacecraft. However, chemicalpropulsion thrusters expend a great deal of mass (propellant) inachieving a predetermined orbit orientation. Electric propulsionthrusters, on the other hand, create significantly less thrust thanthe-chemical propulsion thrusters, but they expend much less mass(propellant) in doing so. That is, electric propulsion thrusters usepropellant (mass) much more efficiently than chemical propulsionthrusters. Using electric and chemical propulsion thrusters to effecttranslation from injection orbit to geosynchronous orbit is described inForte, P. "Benefits of Electric Propulsion for Orbit Injection ofCommunication Spacecraft." American Institute of Aeronautics andAstronautics (AIAA) Paper 92-1955, 14th AIAA International CommunicationSatellite Systems Conference & Exhibit (Mar. 22-26, 1992). A combinedelectric and chemical propulsion system is also described in Free, B."High Altitude Orbit Raising with On-Board Electric Power."International Electric Propulsion Conference Paper 93-205, AmericanInstitute of Aeronautics and Astronautics (AIAA)/AIDA/DGL A/JSASS 23rdInternational Electric Propulsion Conference (Sept. 13-16, 1993).

Because chemical propulsion thrusters exert a much higher force thanelectric propulsion thrusters, they enable translation from injectionorbit to geosynchronous orbit in a substantially shorter period of timethan electric propulsion thrusters. Furthermore, current transfer orbitstrategies for translating a spacecraft from injection orbit togeosynchronous orbit fail to describe a viable burn strategy usingelectric propulsion thrusters exclusively to translate the spacecraft togeosynchronous orbit. Moreover, substitution of electric propulsionthrusters in chemical propulsion thrusters transfer orbit strategieswould require an unacceptable transfer orbit duration (TOD).

Electric propulsion thrusters also introduce yet another consideration,that of stationkeeping and maneuvering. Because electric propulsionthrusters expend substantially less propellant for a given thrust thanchemical propulsion thrusters, and that thrust is relatively lowcompared to chemical propulsion thrusters, they are more desirable forstationkeeping and on-station maneuvers. Because stationkeepingmaneuvers require minimal thrust to reposition the spacecraft, electricpropulsion thrusters perform stationkeeping using much less mass(propellant) than chemical propulsion thrusters.

The trade-off remains that by using chemical propulsion systems andchemical transfer orbit strategies to achieve geosynchronous orbit, asubstantial portion of the spacecraft mass is allocated to propellantfor the chemical thrusters. This mass may be traded into a decreasedrequired launch vehicle capability, or, alternatively, traded forincreased payload for the same launch vehicle capability or acombination thereof. Yet electric propulsion thrusters in combinationwith chemical propulsion transfer orbit strategies provide anunacceptably long transfer orbit duration. Thus, it is desirable toprovide a transfer orbit apparatus and strategy using electricpropulsion thrusters which provides acceptable transfer orbit durationfor a given launch vehicle capability and a given payload.

SUMMARY OF THE INVENTION

The invention provides a spacecraft and method for translating thespacecraft launched into an injection orbit about a central body andoriented in an inertial attitude to a synchronous orbit having asemi-major axis and a predetermined orbital plane. The apparatusincludes a propulsion thruster oriented on the spacecraft to generate athrust having a predetermined force on the spacecraft. The apparatusfurther includes a controller for controlling the timing of firing thepropulsion thruster.

The controller fires the propulsion thruster at apogees of intermediateorbits to successively increase the perigees thereof until thesemi-major axis of an intermediate orbit is substantially equal to thesemi-major axis of the synchronous orbit, thereby defining an initialtransfer orbit for the spacecraft. The controller thereaftercontinuously fires the propulsion thruster to translate the orbit of thespacecraft from the initial transfer orbit to the synchronous orbitwhile maintaining the substantial equality of the synchronous semi-majoraxis and the transfer orbit semi-major axis.

The spacecraft is injected into a particular orbit and its inertialattitude is adjusted to maintain a desired orientation with respect tothe sun. The otherwise unstable injection orbit is selected to offsetthe moments created by adjusting the spacecraft's attitude therebymaintaining a stable orbit. The maintenance of a constant sun angle isparticularly applicable to spinner satellites and 3-axis cubicsatellites in a spin stabilized mode.

The various advantages of the present invention will become apparent toone skilled in the art by reading the following specification and byreference to the following drawing in which:

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a typical injection sequence for launching aspacecraft from ground zero to an injection orbit;

FIG. 2 is a partial exploded view and a partial perspective view of alaunch vehicle;

FIG. 3 depicts a spacecraft as shown in FIG. 2 having mounted thereonelectric propulsion thrusters for effecting translation of thespacecraft to geosynchronous orbit and for performing stationkeepingfunctions;

FIG. 4 illustrates generally the controller portion of the spacecraftfor operating the electric propulsion thrusters for translation togeosynchronous orbit;

FIG. 5 depicts the spacecraft at injection orbit around a central bodyand geosynchronous orbit about the central body;

FIG. 6 depicts the spacecraft in a supersynchronous orbit about acentral body during an apogee burn to raise perigee;

FIG. 7 depicts the angle of inclination between the geosynchronous orbitand an orbit followed by the spacecraft;

FIG. 8 depicts the spacecraft in a series of intermediate orbits about acentral body during a continuous firing sequence;

FIG. 9 depicts the spacecraft having reached geosynchronous orbit;

FIG. 10 is a flow diagram of a sequence implemented by the controllerfor firing the electric propulsion thrusters to reach geosynchronousorbit;

FIG. 11 is a graph of the transfer orbit duration of a spacecraft togeosynchronous orbit versus the weight of the spacecraft delivered togeosynchronous orbit for various transfer orbit propulsion systems;

FIG. 12 depicts a 3-axis cubic satellite with an inertially fixedattitude orbiting around the earth and sun;

FIG. 13 depicts a spinner satellite with an inertially fixed attitudeorbiting around the earth and sun;

FIG. 14 depicts a 3-axis cubic satellite in a spin stabilized mode whoseinertial attitude is synchronized to the sun;

FIG. 15 depicts a spinner satellite whose inertial attitude issynchronized to the sun;

FIGS. 16a and 16b are perspective and plan views, respectively, of aparticular injection orbit;

FIG. 17 is a plot of the node and argument of perigee parameters for theparticular injection orbit; and

FIGS. 18a through 18d depict the injection, transfer and final orbits inwhich the inertial attitude is synchronized to the sun.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The following description of the preferred embodiments is merelyexemplary in nature and not intended to limit the invention or itsapplication or uses. In this specification, note that elements havingsimilar structures or functions will be referred to using like referencenumerals.

Referring to FIG. 1, a spacecraft 10 launched from a position on theearth 12 follows a path predetermined in accordance with parameters ofthe launch vehicle and settles into an elliptical orbit defined as theinjection orbit 16. The optimal transfer orbit trajectory mechanism andapparatus will describe how to translate the spacecraft 10 from theinjection orbit 16 to a a final orbit 18, shown here as a geosynchronousorbit. During the launch sequence, the spacecraft 10 separates fromfirst, intermediate, and final stages as shown at reference numeral 20before reaching injection orbit 16. The number of stages from whichspacecraft 10 separates during launch to injection orbit 16, dependsupon the parameters, capabilities, and particular application for thelaunch vehicle. Such parameters, capabilities, and applications varywidely, but are well known to someone skilled in the art.

Referring to FIG. 2, the spacecraft 10 is shown with other componentparts which comprise launch vehicle 22. Launch vehicle 22 shown in FIG.2 is commonly referred to as a DELTA 7925 launch vehicle and is merelyexemplary of one of a number of launch vehicles well known to oneskilled in the art. The component parts of launch vehicle 22 includethrust augmentation solids 24, first stage oxidizer tank 26, center body28, and fuel tank 30 all of which comprise the first stage of launchvehicle 22. Fuel from fuel tank 30 and reactants from thrustaugmentation solids 24 introduced into first stage oxidizer 26chemically react to yield energy manifested as a thrusting forcepropelling launch vehicle 22. Interstage 32 links fuel tank 30 to secondstage 34 and also provides additional propulsion. Second stage 34includes truss 36 for supporting guidance electronics 38. Guidanceelectronics 38 attaches to spin table 40 which in turn attaches to thirdstage motor 42 via clamp bands 44. Third stage motor 42 attaches tospacecraft 10 via attached fitting 46. During launch, fairings 48provide protection for second stage 34, third stage motor 42, spacecraft10 and the various associated components therewith. After launch vehicleseparation 20 of FIG. 1, only spacecraft 10 continues to injection orbit16.

In operation, launch vehicle 22 is supported on a launch pad (not shown)on the earth 12 and loaded with the appropriate fuels andinstrumentation in preparation for lift-off. At lift-off, first stagecomponents 24-30 operate as described above and provide initial thrustto propel launch vehicle 22 off of the launch pad. At a predeterminedtime in the launch sequence the first stage will detach, at which timeinterstage 32 activates to provide thrust. Similarly, interstage 32,second stage 34, and third stage 42 will provide thrust, then detachfrom the remainder of launch vehicle 22. Also at a predetermined time,the fairings 48 will open, exposing second stage 34, third stage 42, andspacecraft 10. Eventually, after all components have separated fromlaunch vehicle 22, only spacecraft 10 remains and continues to injectionorbit 16. Note that the above described launch vehicle 22 and launchsequence are merely exemplary, and it should be readily recognized byone skilled in the art that the invention described herein is equallyapplicable any of a number of various launch vehicles and launchsequences.

Referring to FIG. 3, a thruster positioning diagram of spacecraft 10 isshown. Spacecraft 10 may be one of any of a number of shapes and sizeswell known in the art, which typically depends upon the particularapplication for which the spacecraft is to be used. Spacecraft 10 has acenter of mass 52 through which passes three axes, a Z axis 54, an Xaxis 58, and a Y axis 60. The Z axis 54, X axis 58, and Y axis 60 areshown extending in a positive direction in accordance with arrows 56,60, and 62, respectively. Spacecraft 50 includes electric propulsionthrusters 50 arranged so that when activated, spacecraft 10 experiencesa force in the positive direction of Z axis 54. Electric propulsionthrusters 50 may optionally be canted so the thrust vector of anyparticular electric propulsion thruster 50 is slightly off parallel fromthe positive Z axis in order to provide redundant directional thrust foruse as a backup, secondary attitude control system. However, theresultant thrust vector when all electric propulsion thrusters firepreferably parallels positive Z axis 54.

Additional electric propulsion thrusters 63 primarily provide thrust forperforming attitude control and stationkeeping of spacecraft 10.Attitude control can also be accomplished with momentum wheels 124,which include gyros with a programmable spin rate. Typically, electricpropulsion thrusters 63 are smaller and provide less thrust thanelectric propulsion thrusters 50 because they only need exert relativelysmall forces to carry out relatively minor spacecraft positioningadjustments. Propulsion thrusters 63 are shown mounted on the same panelas propulsion thrusters 50, and are canted so that individually theyexert thrust inwardly towards center of mass 52. Electric propulsionthrusters 50 are preferably either xenon-ion-propulsion-thrus ters(XIPS) or arcjet propulsion thrusters.

Propulsion thrusters 50 and propulsion thrusters 63 are preferablyelectric propulsion thrusters, and are referred to accordinglythroughout this specification. However, it should be obvious to oneskilled in the art that one or a number of propulsion thrusters 50and/or 63 may be replaced chemical propulsion thrusters, particularly ifthe chemical propulsion thruster has properties similar to electricpropulsion thrusters. As noted above, electric propulsion thrustersprovide much lower thrust than chemical propulsion thrusters, butdeplete significantly less mass for an equivalent period of thrusteractivation.

Referring to FIG. 4, controller 64 selectively activates electricpropulsion thrusters 50 in accordance with a predetermined fire controlroutine. Controller 64 may optionally control the firing of electricthrusters 63 or the spin rate of momentum wheels 124 to provide attitudecontrol. In operation, controller 64 receives input information fromsensor block 65 and determines if electric propulsion thrusters 50should be activated. Sensor block 65 provides input data such asattitude, orientation, and position with respect to earth 12. Controller64 evaluates such information and generates appropriate fire controlsignals responsive to the received information. Controller 64 may alsooptionally control other spacecraft functions and operate as anintegrated spacecraft controller. Furthermore, controller 64 may provideother than simply firing sequences for transfer orbit propulsionthrusters 50. For example, controller 64 may effect communicationbetween spacecraft 10 control systems, may communicate with ground orspace based receivers-transmitters, or may control other instrumentationas required by the particular application.

FIGS. 5-9 depict a transfer orbit trajectory sequence of spacecraft 10from injection orbit 16 to geosynchronous orbit 18. As noted above, likereference numerals will be used to describe similar elements to thosealready previously described. Referring to FIG. 5, spacecraft 10 isshown in injection orbit 16 orbiting around a central body such as theearth 12. Geosynchronous orbit 18 is also depicted in FIG. 5. Injectionorbit 16 is an elliptical orbit having an apogee 66 and a perigee 68which are defined as the highest and lowest points, respectively, of theelliptical orbit with respect to earth 12. It will be recognized by oneskilled in the art that injection orbit 16 is attained in accordancewith a number of predetermined application parameters. Specifically, thespeed of spacecraft 10 at perigee 68 and the height of perigee 68 definethe height of apogee 66 and the eccentricity of tie elliptical injectionorbit 16. Furthermore, to reduce eclipse the orbit's perigee ispreferably positioned behind the earth with respect to the sun.

Referring to FIG. 6, spacecraft 10 is shown in proximity to apogee 66 ofan intermediate orbit 74 having an elliptical shape. When spacecraft 10is oriented in a predetermined attitude and direction using any of anumber of methods known to one skilled in the art, controller 64activates electric propulsion thrusters 50 causing a change in velocityvector ΔV and raising perigee to a new perigee 72 of a new intermediateelliptical orbit 74. The change in velocity vector ΔV substantiallypoints orthogonally to a line between apogee and perigee in thedirection that spacecraft 10 travels when at apogee 66. Firingpropulsion thrusters 50 of spacecraft. 10 during consecutive passesaround apogee 66 raises the perigee 72 of each successive intermediateorbit 74. Intermediate orbit 74 has the same apogee 66, but has anincreased perigee 72 and decreased eccentricity. Note again thatintermediate orbit 74 of FIG. 6 represents a number of intermediateorbits resulting from successive firings of electric propulsionthrusters 50 at apogee 66 to exert a change in velocity ΔV, resulting insuccessive intermediate orbits 74 having successively increasingperigees 72 and decreasing eccentricity. In addition to raising perigee72, firing the electric thrusters 50 around apogee 66 also results in anincrease in the semi-major axis, defined as the average of the apogeeand perigee, of the orbit of spacecraft 10. The semi-major axis israised to be substantially equivalent to the semi-major axis ofgeosynchronous orbit 18, which in accordance with the laws of physics,necessarily implies that the orbital periods of both the geosynchronousorbit 18 (24 hours by definition) and transfer orbit 74 aresubstantially equivalent as well.

Firing electric thrusters 50 at apogee 66 after the spacecraft has beenplaced in the desired burn attitude has two effects. First, as describedabove, firing electrical propulsion thrusters 50 at apogee 66 raisesperigee 72. Second, referring to FIG. 7, firing electric propulsionthrusters 50 at apogee 66 also enables adjustment of the angle ofinclination 73 of the orbital plane 75 of injection orbit 16 withrespect to the geosynchronous plane 77. Because the angle of inclination73 of injection orbit 16 can be no lower in degrees than the latitudefrom which spacecraft 10 was launched, adjustment may be necessary. Inorder to adjust the angle of inclination 73 (to the geosynchronous plane77, for example), it is necessary to incline spacecraft 10 to anattitude so that Z axis 56 does not parallel the plane defined bytransfer orbit 74 and points in the direction of geosynchronous plane77. When this occurs, the angle of inclination 73 of the orbital plane75 of intermediate orbit 74 changes as electric propulsion thrusters 50fire. As described above with respect to intermediate orbit 74, thereare a number of successive intermediate orbital planes 75a-b and anglesof inclinations 73a-b through which spacecraft 10 translates while theangle of inclination 73 is reduced so that orbital plane 75 andgeosynchronous plane 77 substantially coincide.

Firing at such a supersynchronous orbit yields the added benefit thatthe supersynchronicity of the orbit provides a longer burn for raisingperigee 72 and decreasing the angle of inclination 73 than for a lesssupersynchronous orbit. For example, in a subsynchronous orbit, aperigee raising burn may last six out of 10 hours versus nineteen out of22 hours for a supersynchronous orbit. Accordingly, the time of the burnfor rasing perigee 72 lasts longer and raises perigee 72 faster.

Referring to FIG. 7, during the above described process of raisingperigee 72, the angle of inclination 73 between the plane 75 defined bythe orbit followed by the spacecraft (orbit 74 in FIG. 6) and thegeosynchronous plane 77 is adjusted. Spacecraft 10 follows orbit 74, andwhile the direction of thrust primarily points in the direction of thechange in velocity ΔV lying within orbital plane 75, it is also directedout of parallel from orbital plane 75 to sufficiently decrease the angleof inclination 73. Ideally the rate of change of the angle ofinclination 73 and of the increase of perigee 72 is such that the angleof inclination 73 reaches zero degrees and perigee 72 reaches thesemi-major axis of the geosynchronous orbit 16 at substantially the sametime. As described further herein below, when the angle of inclination73 approaches zero degrees, the attitude of the spacecraft is reorientedso that the change in velocity ΔV is substantially parallel to theorbital plane 75 which will then be substantially coincident withgeosynchronous plane 77.

In FIG. 8, after perigee 72 of transfer orbit 74 has been raised to thepredetermined level 78a or orbit 76a, spacecraft 10 oriented so thatwhen controller 64 activates propulsion thrusters 50 a thrust parallelto ΔV as shown in FIG. 8 results along the positive Z axis 54.Spacecraft 10 remains in this predetermined orientation while orbitingaround the earth 12, and also remains in a substantially fixed attitudeso that the positive Z-axis and ΔV are substantially parallel to orbitalplane 76a. Controller 64 fires electric propulsion thrusters 50 aroundthe entirety of intermediate orbits 76a, 76b, and 76c. This continuousfiring has two effects on transfer orbits 76a-c. First, each succeedingtransfer orbit 76a to 76b and 76b to 76c has a progressively higherperigee 78a-c and a progressively lower apogee 66a-c. Second, at thesame time, the eccentricity of the successive orbital ellipses 78a-ddecreases as spacecraft 10 approaches the eccentricity of geosynchronousorbit 18. The 100% duty cycle produces a faster circularization comparedto the much lower duty for GeoStation Transfer Orbit (GTO) sequences.The higher specific impulse of the electric propulsion thrusters meansthat the ΔV inefficiency has a relatively small mass impact. Once ingeosynchronous orbit 18 as shown in FIG. 9, the spacecraft 10 may bereoriented into an operational orientation. For example, as shown inFIG. 9, spacecraft 10 is positioned with positive the Z axis 54 orientedtowards the earth 12.

Referring to FIG. 10, a flow chart depicting the steps carried out bycontroller 62 for translating spacecraft 10 from injection orbit 16 togeosynchronous orbit 18 is shown and will be referred to generally asflow chart 80. At block 82, the launch vehicle 22 of FIG. 2 injectsspacecraft 10 into a supersynchronous orbit 16 as depicted in FIG. 5. Atblock 84, the spacecraft 10 is reoriented to the desired burn attitudeso that when electric propulsion thrusters 50 are activated, theresulting ΔV points in the desired direction. This reorientationmaneuver is typically performed using attitude control propulsionthrusters 62. Control then passes to block 86 where the electricpropulsion thrusters 50 are fired during consecutive passes aroundapogee 66 of transfer orbit 74 of FIG. 6. In block 86, firing propulsionthrusters 50 raises perigee 72 and may also adjust the angle ofinclination 73. Once the orbital plane is properly adjusted as describedwith respect to FIGS. 6 and 7, the spacecraft is reoriented (ifnecessary) at block 88 so that Z axis 54 lies parallel to orbital plane74 and is parallel to ΔV of FIGS. 6.

After reorienting spacecraft 10 to an in-plane attitude, at block 90controller 64 initiates continuous firing of electric propulsionthrusters 50 as spacecraft 10 orbits around earth 12. Such firing isdepicted in FIG. 8, where spacecraft 10 is shown at a number oflocations 10, 10' and 10" following intermediate orbits 76a, 76b, and76c, respectively. This continuous firing results in a continuousraising of perigee 78a-c and a continuous lowering of apogee whilemaintaining a nearly constant semi-major axis for each of intermediateorbits 76a-d which remains substantially equal to the semi-major axis ofgeosynchronous orbit 18. Controller 64 maintains continuous activationof electric thrusters 50 until spacecraft 10 has translated togeosynchronous orbit 18. At block 92, controller 64 tests whethergeosynchronous orbit has been achieved. Once spacecraft 10 achievesgeosynchronous orbit, block 92 passes control to block 94 wherecontroller 64 discontinues firing of electric propulsion thrusters 50.Thereafter, electric propulsion thrusters 50 may be fired as necessaryfor stationkeeping and other maneuvers.

As noted above with respect to FIG. 3, electric propulsion thrusters 50and 62 described herein may be embodied as, but need not be limited to,either XIPS or arcjet electric propulsion systems. Alternatively, asalso noted above, electric propulsion thrusters 50 and 62 may beembodied as a hybrid or complete chemical propulsion system. Referringto FIG. 11, a graph depicts selected performance characteristics of theXIPS and arcjet implementations of electric propulsion thrusters 50 and62 as well as a hybrid of the XIPS and a chemical propulsion system. Inthe graph, the ordinate, depicts the transfer orbit duration (TOD) indays and the abscissa depicts beginning of life weight (BOL) inkilograms. As can be seen from the graph of FIG. 11, while the chemicalpropulsion system has a very short transfer orbit duration, a greatexpense is incurred in terms of the beginning of life weight. Thisimplies that a spacecraft can be put into geosynchronous orbit ratherquickly using chemical propulsion thrusters and chemical transfer orbitstrategies, but with much less instrumentation and stationkeeping fuel.On the other hand, the XIPS and arcjet thrusters demonstrate that whilethe transfer orbit duration is substantially greater than for chemicalpropulsion thrusters, a much greater beginning of life weight isachieved. This translates into substantially more instrumentation andfuel, which may ultimately increase the life expectancy of thespacecraft 50% or greater. Alternatively, for the same spacecraft mass,a launch vehicle having less capability, and a lower cost, can injectthe spacecraft.

Also shown in FIG. 11 is a plot for a hybrid chemical and XIPS system.This system combines features of both chemical and electric propulsionthrusters to both decrease the transfer orbit duration and increase thebeginning of life weight. Along this plot, as the TOD increases, ahigher proportion of electrical propulsion translates the spacecraft togeosynchronous orbit. As the TOD decreases, a lower proportion ofelectrical propulsion translates the spacecraft to geosynchronous orbit.It can be seen from this plot that beginning of life weight can besignificantly increased by use of the hybrid system.

A significant advantage realized by the method and apparatus describedherein is that an equivalent beginning of life weight can be deliveredto geosynchronous orbit 18 using much smaller and less expensive launchvehicles. For example, a spacecraft having a beginning of life weightwhich typically requires launch by an ATLAS or ARIANE 4L booster couldbe delivered by the much less expensive Delta II booster. By the sametoken, greater payloads can be delivered by identical launch vehicles ifelectric propulsion engines translate the spacecraft to geosynchronousorbit. The increased payload can be translated into spacecraft lifetimebecause electric propulsion requires much less mass for stationkeeping,approximately 6 kilograms per year, compared to chemical propulsionsystems, approximately 27 kilograms per year. Further, the time requiredto achieve geosynchronous orbit has been optimized while utilizing asmaller launch vehicle for the same payload.

FIG. 12 depicts a 3-axis cubic satellite 102 orbiting the earth 100,which in turn orbits the sun 104 for the orbit transfer sequence shownin FIGS. 5-9. The satellite 102 is injected into a naturally stableorbit with perigee positioned behind the earth. This reduces theduration of the eclipse caused by the earth coming between the sun andthe satellite. The satellite has a fixed inertial attitude in line withthe velocity vector at apogee so that thrust at apogee increases perigeeand thrust at perigee reduces apogee. As the earth moves around the sun,the orientation between the satellite and the sun changes atapproximately 1° per day. Thus, to maintain a high and uniform powerlevel, the satellite's solar panels 106 are rotated so that they alwaysface the sun.

A spinner 108 is typically a cylindrical satellite having solar panels110 affixed to its outer surface, but not to either of its ends. Thespinner 108 spins around its central axis 110, and thus is spinstabilized. It is well known in the art that spin stabilization reducesthe complexity of maintaining an inertial attitude. Unlike the 3-axiscubic satellite, the spinner's solar panels are fixed and thus can notbe independently rotated to follow the sun. Furthermore, the spinner'ssolar panel surface area at a given orientation to the sun is much lessthan that of a 3-axis cubic satellite. As shown in FIG. 13, as the earthmoves around the sun, the orientation of the spiriner's solar panels tothe sun would change by about 1° per day, and consequently the powerproduced by the panels would be reduced. After approximately 20 days,the power produced would be insufficient to power the satellite. Becausethe transfer orbits associated with electrical propulsion typically takebetween 30 and 90 days, this strategy can not be used with spinners. Asa result, spinners are currently transferred to their final orbits usingchemical propulsion.

An alternate transfer orbit sequence, as shown in FIGS. 14 and 15,allows the 3-axis cubic satellite to be spin stabilized during transferand allows spinners to be transferred with an electrical propulsionsystem. A number of 3-axis cubic and spinner satellites are described inthe Hughes Space and Communications Group Brochure, 1989. To accomplishthis, the transfer orbit sequence described in FIGS. 5-9 is modified intwo ways. First, the inertial attitude 112 of the satellite is adjustedto maintain a desired orientation with respect to the sun so that itssolar panels face the sun. Preferably 0° for the cubic satellite and 90°for the spinner. This allows the solar panels on the 3-axis cubicsatellite to be deployed in a fixed position, in which the satellite canbe spin stabilized. Adjusting the panels while spinning the satellitewould create an immense torque that would damage the satellite.Furthermore, adjusting the inertial attitude creates a turning momentthat tends to destabilize the transfer orbit. Second, the satellite isinjected into a particular orbit 114 to create a moment that offsets thedestabilization effects of adjusting the inertial attitude. Together theattitude adjustment and orbit selection provide a stable orbit and asun, as well as earth, synchronized inertial attitude.

FIGS. 16a and 16b are perspective and plan views of the injection orbit114 about the earth 100. In general, an orbit is completely defined byfive parameters: semi-major axis (sma), eccentricity (ecc), inclination(i), right ascension of ascending node (Ω) and argument of perigee (ω).The sma is the average of the apogee (A) and perigee (P). The ecc is thedifference of the apogee and perigee divided by their sum, and is ameasure of the "roundness" of the orbit. The inclination is the anglebetween the plane of the orbit and a reference plane. The node is theangle from a reference vector (γ), commonly known as Aries (vernalequinox), in the Earth Centered Inertial (ECI) coordinate system, to thepoint where the ascending orbit intersects the reference plane. Theargument of perigee is the angle from the ascending node to perigee.

As discussed above, the sma, ecc and i parameters are selected inaccordance with the launch vehicle, payload and the desired final orbit.The node and argument of perigee parameters are selected so that theinjection and transfer orbits create moments that substantially offset amoment created by adjusting the satellite's inertial attitude. As aresult, the satellite maintains a stable orbit while remainingsynchronized to the sun.

FIG. 17 is a plot 116 of the node (Ω) and argument of perigee (ω) over atime of delivery (TOD) to the final orbit for a given inertial attitudeadjustment. The dashed lines indicate (ω, Ω) pairs which create momentsthat do not balance the moment created by the attitude adjustment. Thisdestabilizes the orbit, which changes the values for (ω, Ω) over time.The solid line indicates the (ω,Ω) pair that stabilizes the orbit andthus, maintains substantially constant (ω, Ω) values over the ToD. Inthe inertially fixed attitude transfer strategy, an argument of perigeeof either 0° or 180° would provide a naturally stable orbit. In theadjusted inertial attitude strategy, the argument of perigee is offsetfrom 0° or 180° to cancel the other moment.

A simulation of the satellite's orbit is used to generate plot 116.Although complex, simulations that incorporate the relative motions ofthe sun, earth, and orbiting satellite are well known. The knownparameters such as the initial apogee, perigee, eccentricity, andinclination, the sma of the final orbit, and the adjustment rate of theinertial attitude, are specified. The (ω, Ω) pair is then determined byfirst selecting the desired sun angle, e.g. preferably 0° for a 3-axiscubic and 90° for a spinner. The desired sun angle and (γ) determine thevalue of the node. As a result, the argument of perigee is the onlyremaining variable. A value of the argument of perigee is selected andcorrected until both the node and the argument of perigee remainconstant over the ToD.

Although the desired sun angle is preferably 0° for a 3-axis cubic and90° for a spinner, most applications will allow variation from thepreferred angle within some tolerance. This reduces power efficiency anduniformity but increases the stability of the orbit. As a result, the(ω, Ω) pair become less sensitive and thus easier to select.

FIGS. 18a through 18d illustrate the alternate orbit transfer strategyfrom injection to final orbit for the 3-axis cubic satellite 102. Thissame strategy applies equally to the spinner satellite. Referring toFIG. 18a, satellite 102 is shown in injection orbit 114 around the earth100 with its solar panels 106 deployed in a fixed position. The 3-axiscubic satellite preferably spins around its central axis 117 whichpoints towards the sun 104.

As shown, injection orbit 114 has a particular (ω, Ω) pair selected tomaintain orbit stability throughout the transfer orbit. Referring toFIG. 18b, when satellite 102 is at apogee 118, the controller activatesthe electric propulsion thruster causing a change in velocity vector ΔVand raising perigee 120 to a new perigee 120 of a new intermediateelliptical orbit 114. This raises the semi-major axis to besubstantially equivalent to the semi-major axis of the final orbit 122so that the orbital periods of the transfer and final orbits aresubstantially equivalent as well. As shown the final orbit 122 iscircular, e.g., geo, meo or leo synchronous.

As described above in FIG. 7, firing the electrical propulsion thrustersat apogee also enables adjustment of the angle of inclination of theorbital plane of the injection orbit 114 with respect to the referenceplane of the earth. Ideally the rate of change of the angle ofinclination i and of the increase of perigee 120 is such that the angleof inclination i reaches zero degrees and the semi-major axis of theintermediate orbit reaches the semi-major axis of the final orbit 122 atsubstantially the same time.

In FIG. 18c, after perigee 122 of transfer orbit 114 has been raised tothe predetermined level, the satellite is oriented in the final orbitalplane with a desired inertial attitude 112 so that when the controlleractivates the propulsion thruster a thrust component parallel to ΔVresults. Satellite 102 remains in the final orbital plane while orbitingaround the earth 100, and adjusts its inertial attitude so that solarpanels 106 maintain a desired angle to the sun 104. The satellite'sinertial attitude is preferably adjusted in one of two ways. First, thecontroller fires the electric propulsion thrusters 50 shown in FIG. 3 toproduce an approximately instantaneous unbalanced thrust at the samepoint in successive orbits. For example, the thrusters can be fired atsuccessive apogees to produce a 1° per day adjustment. Second, themomentum wheels 124 shown in FIG. 3 can be programmed to provide aconstant inertial adjustment equivalent to 1° per day.

While orbiting the earth, the controller fires electric propulsionthrusters around the entirety of the intermediate orbits 114. Thiscontinuous firing has two effects on the transfer orbits 114. First,each succeeding transfer orbit has a progressively higher perigee 120and a progressively lower apogee 118. Second, at the same time, theeccentricity of the successive orbital ellipses decreases as satellite102 approaches the eccentricity of the final orbit 114. Once ingeosynchronous orbit 122 as shown in FIG. 18d, the satellite 102 may bereoriented into an operational orientation. For example, as shown inFIG. 18d, satellite 102 is positioned with its central axis 117 orientedtowards the earth 100.

Although the invention has been described with particular reference tocertain preferred embodiments thereof, variations and modifications canbe effected within the spirit and scope of the following claims. Forexample, the orbit transfer strategies can be applied to ellipticalfinal orbits as well as circular orbits.

I claim:
 1. A method for launching a spacecraft into a final synchronousorbit about a central body which orbits about a sun, said final orbithaving a semi-major axis and an eccentricity, comprising:injecting saidspacecraft into an initial elliptical orbit about the central body andhaving an inertial attitude, said injection orbit being inclined withrespect to a reference plane of said central body and being defined byan apogee, a perigee, a node that is the angle measured from a referencevector to the point where the ascending orbit intersects the referenceplane, and an argument of perigee that is the angle from the node toperigee; adjusting the perigees of successive intermediate orbits untiltheir semi-major axis is substantially equal to the semi-major axis ofthe final orbit; continuously providing thrust to change theeccentricity of the intermediate orbit to that of said final orbit whilemaintaining the substantial equality of the final semi-major axis andthe intermediate orbit semi-major axis; and adjusting the inertialattitude of the spacecraft to maintain, within a tolerance, a selectedorientation of the spacecraft to the sun, said node and argument ofperigee being selected to create an orbit moment that opposes a momentproduced by adjusting the inertial attitude so that said node andargument of perigee remain substantially constant over a time ofdelivery to the final orbit.
 2. The method of claim 1, wherein saidspacecraft has a central axis that is oriented with said inertialattitude and has solar panels, the method further comprising:spinningthe spacecraft about its central axis thus reducing the complexity ofmaintaining and adjusting the inertial attitude, which maintains anorientation of the spacecraft's solar panels to the sun to provide anincreased and approximately uniform power level.
 3. The method of claim2, wherein said spacecraft is a spinner satellite in which said solarpanels are affixed to its outer surface directed away from its centralaxis and spin around its central axis.
 4. The method of claim 3, whereinsaid spinner satellite's central axis is maintained at an approximatelyperpendicular orientation to a radius between the sun and the satellite.5. The method of claim 4, wherein said spacecraft is a 3-axis cubicsatellite in which said solar panels are deployed in a fixed positiondirected towards the sun.
 6. The method of claim 5, wherein said 3-axiscubic satellite is maintained approximately in line with a radiusbetween the sun and the satellite.
 7. The method of claim 2, wherein thespacecraft includes an electrical propulsion thruster for providing thecontinuous thrust.
 8. The method of claim 7, wherein said electricalpropulsion thruster provides art unbalanced thrust pulse duringsuccessive orbits to adjust the inertial attitude.
 9. The method ofclaim of claim 7, wherein the spacecraft includes a momentum wheel thatcontinuously adjusts the inertial attitude of the spacecraft.
 10. Amethod for launching a spacecraft into a circular orbit about a centralbody which orbits about a sun, said circular orbit having a semi-majoraxis, comprising:providing a satellite having a central axis, saidsatellite include solar panels in a known orientation to said centralaxis, an electrical propulsion thruster for providing thrust along saidcentral axis, and an attitude controller for adjusting the inertialattitude of the satellite's central axis injecting said satellite intoan initial supersynchronous elliptical orbit about the central body andhaving an inertial attitude that orients the solar panels towards thesun, said injection orbit being inclined with respect to a referenceplane of said central body and being defined by an apogee, a perigee, anode that is the angle measured from a reference vector to the pointwhere the ascending orbit intersects the reference plane, and anargument of perigee that is the angle from the node to perigee; spinningthe satellite about its central axis to reduce the complexity ofmaintaining its inertial attitude; firing the electrical propulsionthruster at apogees of intermediate orbits to successively increase theperigees thereof until their semi-major axis is substantially equal tothe semi-major axis of the circular orbit; continuously firing theelectrical propulsion thruster to translate the orbit of the satellitefrom the intermediate orbit to the circular orbit while maintaining thesubstantial equality of the circular semi-major axis and theintermediate orbit semi-major axis; and controlling the attitudecontroller to adjust the inertial attitude of the spacecraft tomaintain, within a tolerance, a selected orientation of the spacecraft'ssolar panels to the sun, to provide an increased and approximatelyuniform power level; said node and argument of perigee being selected tocreate an orbit moment that opposes a moment produced by adjusting theinertial attitude so that said node and argument of perigee remainsubstantially constant over a time of delivery to the circular orbit.11. The method of claim 10, wherein said satellite is a spinner in whichsaid solar panels are affixed to its outer surface directed away fromits central axis and spin around its central axis.
 12. The method ofclaim 11, wherein said spinner satellite's central axis is maintained atan approximately perpendicular orientation to a radius between the sunand the satellite so that its solar panels face the sun.
 13. The methodof claim 10, wherein said satellite is a 3-axis cubic satellite in whichsaid solar panels are deployed in a fixed position directed towards thesun.
 14. The method of claim 13, wherein said 3-axis cubic satellite ismaintained approximately in line with a radius between the sun and thesatellite.
 15. The method of claim 10, wherein said attitude controllercomprises the electrical propulsion thruster which provides anunbalanced thrust pulse during successive orbits to adjust the inertialattitude.
 16. The method of claim of claim 10, wherein the attitudecontroller comprises a momentum wheel that continuously adjusts theinertial attitude of the spacecraft.